WO2022189154A1

WO2022189154A1 - Fuel conditioning system and method configured to supply an aircraft turbine engine using fuel from a cryogenic tank

Info

Publication number: WO2022189154A1
Application number: PCT/EP2022/054594
Authority: WO
Inventors: Pierre-Alain LAMBERT; Samer MAALOUF
Original assignee: Safran
Priority date: 2021-03-08
Filing date: 2022-02-24
Publication date: 2022-09-15
Also published as: FR3120393A1; EP4305287A1; US20240043135A1; CN116917608A

Abstract

Disclosed is a fuel conditioning system (SC) configured to supply an aircraft turbine engine using fuel (Q) from a cryogenic tank (RC), the conditioning system (SC) comprising at least one pumping turbomachine (1), a first heat exchanger (31) configured to heat the fuel (Q) in the fuel circuit (CQ) by circulating an air stream (A), and at least one heating turbomachine (2) configured to supply an air flow (A) to the first heat exchanger (31), the heating turbomachine (2) comprising an air intake compressor (21), a combustion chamber (24) and an air exhaust turbine (22) configured to drive the air intake compressor (21), the combustion chamber (24) being supplied with air taken from the air flow (A) and with the fuel (Q) from the fuel circuit (CQ).

Description

Fuel conditioning system and method configured to supply an aircraft turbine engine with fuel from a cryogenic tank

The present invention relates to the field of aircraft comprising turbine engines powered by fuel stored in a cryogenic tank.

It is known to store fuel, in particular hydrogen, in liquid form to limit the size and weight of aircraft tanks. For example, the fuel is stored at a temperature of the order of 20 to 22 Kelvins (-253 to -251°C) in a cryogenic tank of the aircraft.

In order to be injected into the combustion chamber of a turbine engine, the fuel must be pumped and heated in order to allow optimal combustion. Such a heating step is for example necessary to reduce the risk of icing of the water vapor contained in the air which circulates in the turbine engine, in particular, at the level of the fuel injectors of the turbine engine.

In practice, the fuel heating step is energy-intensive and requires taking calories from hot sources of the aircraft. For example, the heat generated by the turbine engine can be used (heat from the lubricating oil, calories at the turbine outlet, heat from the nozzle, etc.). Heat from the aircraft can also be used (air from the cabin, heat from electrical or electronic systems, etc.).

One of the issues is to optimize the heating of the fuel while taking advantage of the cooling properties of the fuel during its transport from the cryogenic tank to the combustion chamber of the turbine engine while guaranteeing high safety.

In the prior art, with reference to FIGS. 1 and 2, several architectures have been proposed for conducting fuel Q from a cryogenic tank RC to the combustion chamber CC of a turbine engine T. In known manner, the cryogenic tank RC belongs to the aircraft reference frame REF-A while the combustion chamber CC belongs to the turbine engine reference frame REF-T.

Hereafter, the terms "upstream" and "downstream" are defined with respect to the flow direction of the fuel Q from the cryogenic tank RC to the combustion chamber CC.

With reference to the representing a first architecture, it has been proposed to successively provide, from upstream to downstream, a pump 101, belonging to the reference of the turbine engine REF-T, and a heat exchanger 102 taking calories from the turbine engine T, in particular, at the level of the nozzle of the turbine engine T before injecting fuel Q into the combustion chamber CC. The pump 101 is connected to the cryogenic tank RC by a transport pipe 100 which provides the interface between the two repositories REF-A, REF-T.

Such a first architecture has several drawbacks. First of all, the fuel Q is led with a low pressure and a low temperature in the transport line 100 to the pump 101 which must therefore be thermally insulated, which has disadvantages. Furthermore, the presence of a heat exchanger 102 in the nozzle of the turbine engine T impacts the performance of the turbine engine T. 'significant thermo-fluidic instabilities, which imposes long cooling down or re-cooling times that are detrimental to the safety and operability of the turbine engine T.

With reference to the representing a second architecture, it has been proposed to deport the pump 101 as close as possible to the cryogenic tank RC, that is to say in the frame of reference of the aircraft REF-A in order to reduce the size in the frame of reference of the turbine engine REF-T. The fuel Q is compressed with a high pressure and a low temperature in a transport pipe 100' connecting the pump 101 to the exchanger 102. Similarly to above, the transport pipe 100' must be thermally insulated but also mechanically reinforced. to withstand the high fuel pressure Q.

In both architectures, due to the very low temperature of the fuel Q in the transport lines 100, 100', only helium can be used as a purge or conditioning gas in the event of a fuel Q leak. Helium should be avoided for commercial aeronautical application since this fluid is a high cost noble gas. Known in the prior art by the patent application EP3623604A1 a military propulsion turbomachine intended to operate at very high speed which comprises a heat exchanger configured to take calories from the exhaust so as to supply them to a flow of fuel.

The invention thus aims to eliminate at least some of these drawbacks by proposing a new fuel conditioning system.

PRESENTATION OF THE INVENTION

au moins une turbomachine de pompage, comprenant une pompe et une turbine configurée pour entrainer la pompe, ladite pompe étant configurée pour prélever du carburant issu du réservoir cryogénique et le faire circuler d’amont en aval dans un circuit de carburant comprenant une sortie d’alimentation du turbomoteur, la turbine étant montée dans le circuit de carburant de manière à permettre l’entrainement en rotation de la turbine par la circulation du carburant,
un premier échangeur thermique, monté en amont de la turbine, configuré pour réchauffer le carburant dans le circuit de carburant par circulation d’un flux d’air,
au moins une turbomachine de chauffage configurée pour alimenter le premier échangeur thermique avec un flux d’air, la turbomachine de chauffage comportant un compresseur d’admission d’air, une chambre de combustion et une turbine d’échappement d’air configurée pour entrainer le compresseur d’admission d’air, la chambre de combustion étant alimentée par de l’air prélevé dans le flux d’air et par du carburant issu du circuit de carburant.

The invention relates to a fuel conditioning system configured to supply an aircraft turbine engine using fuel from a cryogenic tank, the conditioning system comprising:

at least one pumping turbomachine, comprising a pump and a turbine configured to drive the pump, said pump being configured to take fuel from the cryogenic tank and circulate it from upstream to downstream in a fuel circuit comprising an outlet from powering the turbine engine, the turbine being mounted in the fuel circuit so as to allow the rotational drive of the turbine by the circulation of the fuel,
a first heat exchanger, mounted upstream of the turbine, configured to heat the fuel in the fuel circuit by circulation of an air flow,
at least one heating turbine engine configured to supply the first heat exchanger with an air flow, the heating turbine engine comprising an air intake compressor, a combustion chamber and an air exhaust turbine configured to drive the air intake compressor, the combustion chamber being supplied with air taken from the air flow and with fuel from the fuel circuit.

Advantageously, the use of a pumping turbomachine coupled to a heating turbomachine makes it possible to heat the fuel autonomously while consuming a small fraction of fuel. The fuel supplied at the outlet has a temperature compatible with the turbine engine, preferably a temperature close to ambient temperature and sufficient pressure to be able to be led to the turbine engine. Thus, the conditioning system can advantageously be positioned in the frame of reference of the aircraft and not in that of the turbine engine, preferably close to the cryogenic tank. Advantageously, nitrogen can be used as neutral gas for purges and scans, the implementation of the conditioning system is greatly facilitated.

The use of a first heat exchanger makes it possible to transmit the calories of the heating turbomachine to the fuel in order to heat it optimally. Advantageously, the pumping turbomachine is driven by the circulation of fuel in the fuel circuit, which makes it possible to simplify the design of the pumping turbomachine by allowing low fuel leakage between the pump and the turbine, and increases the security.

Preferably, the combustion chamber is fed by air taken from the air flow and by fuel from the fuel circuit and taken downstream of the turbine. The fuel advantageously has an optimum temperature and pressure for the combustion chamber.

Preferably, the system comprises a second heat exchanger configured to heat the fuel in the fuel circuit by circulation of an intake air flow from the heating turbomachine. Thus, the intake air, which circulates at a lower temperature than the exhaust air, makes it possible to preheat the fuel before it is reheated by the first heat exchanger. Such preheating makes it possible on the one hand to reduce the risk of condensation of the hot gases passing through the first exchanger, and on the other hand to reduce the consumption of the air intake compressor of the heating turbomachine.

According to one aspect of the invention, the system comprises an auxiliary heat exchanger configured to heat the fuel in an upstream portion of the fuel circuit by circulating fuel flowing downstream of the turbine. Advantageously, an exchange of calories by circulation of fuel having different temperatures makes it possible to carry out a gradual preheating. Such preheating reduces the risk of air condensation in the following exchangers.

According to one aspect of the invention, the system comprises an electric generator configured to be driven by the heating turbomachine. According to one aspect, the electric generator is driven by a free turbine configured to be driven by an air flow from the heating turbomachine. Thus, the pumping turbomachine makes it possible to generate an air flow for the first heat exchanger which is used by the auxiliary turbine to produce electricity. The conditioning system thus produces electricity to contribute to its autonomy.

According to another aspect, the electric generator is driven directly by the heating turbomachine or via a gear train.

Preferably, the heating turbomachine is configured to implement a Brayton cycle so as to generate significant heat by combustion of a fraction of the fuel.

Preferably, the pumping turbomachine is entirely bathed in fuel. Thus, it is not necessary to ensure a strong seal between the pump and the turbine. Advantageously, guidance by fluid steps can be implemented in a practical way.

Preferably, the pumping turbomachine is a hydrogen expansion turbopump in order to allow it to be driven by the circulation of the fuel when the latter is hydrogen.

The invention also relates to a set of at least one cryogenic tank, an aircraft turbine engine and a conditioning system, as presented previously, fluidly connecting the cryogenic tank and the aircraft turbine engine.

Preferably, the conditioning system is positioned close to the cryogenic tank and connected to the turbine engine by at least one circulation pipe which circulates fuel at ambient temperature. Such a pipe advantageously has a simple structure which does not require thermal insulation or special cooling.

According to one aspect, the conditioning system is supplied with air by an air flow external to the turbine engine. According to another aspect, the conditioning system is supplied with air by an air flow from the turbine engine, in particular, from a compressor of the turbine engine.

Preferably, the assembly comprises at least one fuel buffer storage capacity configured to be supplied by the conditioning system and configured to supply the aircraft turbine engine.

Preferably, the fuel buffer storage capacity is configured to store fuel at high pressure and at ambient temperature. Thus, the fuel of a buffer storage capacity can be used directly by a turbine engine and, in a reactive way, according to the needs of the turbine engine.

la pompe de la turbomachine de pompage prélève du carburant dans le réservoir cryogénique et le fait circuler d’amont en aval dans un circuit de carburant,
le carburant est réchauffé dans le circuit de carburant par le premier échangeur thermique par circulation d’un flux d’air issu de la turbomachine de chauffage,
la circulation du carburant dans le circuit de carburant entrainant en rotation la turbine de la turbomachine de pompage,
la chambre de combustion de la turbomachine de chauffage étant alimentée par de l’air prélevé dans le flux d’air et par du carburant issu du circuit de carburant.

The invention also relates to a process for conditioning fuel by means of a conditioning system, as presented above, for supplying an aircraft turbine engine from fuel coming from a cryogenic tank, process in which:

the pump of the pumping turbomachine takes fuel from the cryogenic tank and circulates it from upstream to downstream in a fuel circuit,
the fuel is heated in the fuel circuit by the first heat exchanger by circulation of an air flow coming from the heating turbomachine,
the circulation of the fuel in the fuel circuit driving the turbine of the pumping turbine engine in rotation,
the combustion chamber of the heating turbomachine being supplied with air taken from the air flow and with fuel from the fuel circuit.

PRESENTATION OF FIGURES

The invention will be better understood on reading the following description, given by way of example, and referring to the following figures, given by way of non-limiting examples, in which identical references are given to similar objects. .

The is a schematic representation of a first architecture for conditioning fuel from a cryogenic tank to a turbine engine according to the prior art.

The is a schematic representation of a second architecture for conditioning fuel from a cryogenic tank to a turbine engine according to the prior art.

The is a schematic representation of an architecture for conditioning fuel from a cryogenic tank to a turbine engine with a conditioning system according to the invention.

The is a schematic representation of a first embodiment of the conditioning system.

The is a schematic representation of a second embodiment of the conditioning system.

The is a schematic representation of a third embodiment of the conditioning system.

The is a schematic representation of a conditioning system. to supply a buffer storage capacity with fuel.

It should be noted that the figures expose the invention in detail to implement the invention, said figures can of course be used to better define the invention if necessary.

DETAILED DESCRIPTION OF THE INVENTION

With reference to the , there is shown an architecture according to one embodiment of the invention for conducting fuel Q from a cryogenic tank RC to the combustion chamber CC of a turbine engine T of an aircraft. In a known manner, the cryogenic tank RC belongs to the reference frame of the aircraft REF-A while the combustion chamber CC belongs to the reference frame of the turbine engine REF-T.

In this example, the fuel is liquid hydrogen but the invention applies to other types of fuel, for example, liquid methane or liquefied natural gas.

According to the invention, there is provided a fuel conditioning system SC configured to supply the combustion chamber CC of the turbine engine T from fuel in the liquid phase coming from the cryogenic tank RC. The conditioning system SC belongs to the aircraft standard REF-A and makes it possible to pump the fuel Q as close as possible to the cryogenic tank RC. The space requirement in the REF-T turbine engine repository is thus reduced. As will be presented later, the conditioning system SC makes it possible to heat the fuel to an optimum temperature by circulation of an air flow from an air inlet EA. The fuel conditioning system SC is connected to the turbine engine T by a circulation line 6.

With reference to the , the conditioning system SC comprises a pumping turbomachine 1, a fuel circuit CQ, a heating turbomachine 2, a first heat exchanger 31 and a second heat exchanger 32.

As shown in , there is shown a pumping turbomachine 1 (also called a turbopump), comprising a pump 11 and a turbine 12 configured to drive the pump 11, the pump 11 being configured to take fuel Q from the cryogenic tank RC and circulate it from upstream to downstream in the CQ fuel system. In this example, the pump 11 and the turbine 12 are integrally connected in rotation by a shaft 13.

The CQ fuel system (in close dashes on the ) thus comprises an inlet configured to be fluidically connected to the cryogenic tank RC and a supply outlet S of the turbine engine T.

The turbine 12 is mounted in the fuel circuit CQ so as to be driven in rotation by the circulation of the fuel Q. In other words, the pump 11 makes it possible to increase the pressure of fuel Q in the fuel circuit CQ downstream of the pump 11. This pressure drives the turbine 12.

The pumping turbomachine 1 is preferably in the form of a hydrogen expansion turbopump. Advantageously, the turbine 12 and the pump 11 are entirely bathed in fuel, in particular pure hydrogen, which reduces the problems of dynamic sealing linked to the drive, in particular, as may be the case with an electric motor drive. Any leak at pump 11 can be reinjected into turbine 12, upstream or downstream of the latter. The rotational guidance can advantageously be ensured by fluid bearings guaranteeing an optimum service life without the need for lubrication.

As shown in , there is shown a heating turbine engine 2 configured to take a flow of air A from an air inlet EA, preferably outside the aircraft or in the turbine engine T, and to supply the first heat exchanger 31 By way of example, the air inlet EA may correspond to an air bleed from a compressor of the turbine engine T.

The heating turbomachine 2 comprises an air intake compressor 21, a combustion chamber 24 and an air exhaust turbine 22 configured to drive the air intake compressor 21.

In this example, the air intake compressor 21 and the air exhaust turbine 22 are integrally connected in rotation by a shaft 23. The heating turbomachine 2 makes it possible to take an air flow A upstream for evacuate it downstream after heating.

In this embodiment, the first heat exchanger 31 is supplied with an exhaust air flow A downstream (at high temperature) while the second heat exchanger 32 is supplied with an air flow A upstream from the exhaust. intake (at a lower temperature than the exhaust). Preferably, the second heat exchanger 32 is positioned upstream of the first heat exchanger 31 on the fuel circuit CQ. Thus, the second heat exchanger 32 performs a preheating function with respect to the first heat exchanger 31 which provides the main heating. Such an arrangement makes it possible to optimize the efficiency of the heating and to avoid condensation and/or icing of the water contained in the air in contact with the fuel Q in the first heat exchanger 31. Thus, the output S of the circuit CQ fuel tank is supplied with fuel at optimum temperature.

According to the invention, the combustion chamber 24 is supplied with air taken from the air flow A and with fuel Q from the fuel circuit CQ and taken downstream of the turbine 12. Thus, the combustion chamber combustion 24 of the heating turbomachine 2 is supplied with fuel Q at optimum temperature and at optimum pressure.

The heating turbomachine 2 is preferably in the form of a Brayton cycle turbomachine and makes it possible to generate a flow of hot air so as to provide calories to the fuel Q.

Still with reference to the , the conditioning system SC comprises a bifurcation member 4, comprising for example a 3-way valve, configured to receive fuel Q from the turbine 12 as an input and to supply, on the one hand, the combustion chamber 24 of the heating turbine engine 2 and, on the other hand, the supply output S of the turbine engine T.

According to an optional aspect, with reference to the , the conditioning system SC further comprises an auxiliary heat exchanger 33 configured to heat the fuel Q in an upstream portion of the fuel circuit CQ by circulation of fuel Q flowing downstream of the turbine 12. Preferably, the heat exchanger auxiliary 33 is positioned close to the pumping turbomachine 1. The auxiliary heat exchanger 33 of the fuel/fuel type makes it possible to gradually heat the fuel at the outlet of the cryogenic tank RC. This makes it possible to avoid condensation and/or icing of the water contained in the air in contact with the fuel Q in the second heat exchanger 32 and/or the first heat exchanger 31.

It goes without saying that at least one heat exchanger could be positioned upstream of the turbine 12 or upstream of the first heat exchanger 31 to provide calories from heat sources from the aircraft and/or from the turbine engine T.

According to an optional aspect, with reference to the , the conditioning system SC further comprises an electric generator 51 driven by a free turbine 52 configured to be driven by a flow of air A coming from the heating turbomachine 2. In this example, the electric generator 51 is driven by the free turbine 52 via a shaft 53. The rotation of the free turbine 52 advantageously makes it possible to generate electrical energy to supply electrical energy, for example, to the non-propulsive functions of the aircraft, which makes it possible to limit the mechanical withdrawals on the turbine engine.

According to another aspect of the invention, the electric generator 51 can be coupled to the heating turbomachine 2, in particular, directly or via a gear train (not shown).

A fuel conditioning method using an SC conditioning system according to the invention will now be presented with reference to the .

During the process, the pump 11 of the pumping turbomachine 1 takes fuel from the cryogenic tank RC and circulates it from upstream to downstream in a fuel circuit CQ. When pumping, the fuel pressure Q rises.

Advantageously, the circulation of the fuel Q in the fuel circuit CQ drives the turbine 12 of the pumping turbomachine 1 in rotation, which makes it possible to use the enthalpy of the fuel Q to generate a movement. In addition, following training, the fuel pressure Q is optimal for supplying a combustion chamber, namely, both that of the turbine engine T and that of the heating turbine engine 2.

In this example, the fuel Q is first of all heated in the fuel circuit CQ by the second heat exchanger 32 by circulation of an upstream air flow A drawn off by the heating turbomachine 2. This air flow A provides preheating. Then, the fuel Q is heated in the fuel circuit CQ by the first heat exchanger 31 by circulation of a flow of air A downstream from the heating turbine engine 2. The combustion chamber 24 of the heating turbine engine 2 is fed by air taken from the air flow A and by fuel Q from the fuel circuit CQ and taken downstream of the turbine 12. The heating turbomachine 2 makes it possible to implement a Brayton cycle to bring calories directly to the CQ fuel system. The temperature and the pressure of the fuel Q are thus increased by the conditioning system SC to directly supply the turbine engine T.

With reference to the , the positioning of the conditioning system SC in the aircraft frame of reference REF-A makes it possible to reduce the size in the frame of reference of the turbine engine REF-T. Such a conditioning system SC makes it possible to ensure a compromise between simplicity and safety of implementation for the circulation pipe 6 connecting the conditioning system SC to the turbine engine T. Indeed, as the conditioning system SC makes it possible to increase the pressure and to increase the temperature of the fuel Q, it is no longer necessary to isolate the circulation line 6 connecting the conditioning system SC to the turbine engine T. The fuel Q is compressed and heated up to a supercritical state where its implementation no longer presents any difficulty with respect to the thermofluidic instabilities inherent in cooling. The usual technologies for the distribution of gas under pressure can then advantageously be implemented instead of more complex solutions required by cryogenic fluids.

Advantageously, thanks to the conditioning system SC, the conditioning of the fuel in a cryogenic tank RC and the use of the fuel in the turbine engine T are decoupled, which simplifies management. Advantageously, the speeds of rotation of the pumping turbomachine 1 and of the heating turbomachine 2 can be adjusted independently in order to optimize the heating of the fuel Q.

In addition, it is advantageously no longer necessary to place an exchanger at the level of the nozzle of the turbine engine T as in the prior art. The performance of the turbine engine T, in particular related to the flow of the primary flow, is not affected.

Advantageously, nitrogen can be used as inert gas for the purges and sweeps of the circulation line 6, which facilitates the implementation compared to helium. Finally, the conditioning system SC makes it possible to operate autonomously by consuming a fraction of the fuel Q pumped into the cryogenic tank RC.

According to another aspect of the invention, with reference to the , the conditioning system SC is configured to supply one or more fuel buffer storage capacities 7, hereinafter storage capacities 7. Such storage capacities 7 make it possible to ensure a sufficient flow rate of fuel Q during the transient phases of the turbine engine T or when the conditioning system SC does not provide sufficient flow. When the flow rate fuel demand is lower, the conditioning system SC supplies the storage capacities 7. Thus, even in degraded conditions, the turbine engine T has in all circumstances a fuel at high pressure and at ambient temperature ready to be used by the turbine engine T.

Claims

Fuel conditioning system (SC) configured to supply an aircraft turbine engine (T) from fuel (Q) coming from a cryogenic tank (RC), the conditioning system (SC) comprising:

at least one pumping turbine engine (1), distinct from the aircraft turbine engine (T), comprising a pump (11) and a turbine (12) configured to drive the pump (11), the pump (11) being configured to take fuel (Q) from the cryogenic tank (RC) and circulating it from upstream to downstream in a fuel circuit (CQ) comprising a supply outlet (S) of the turbine engine (T), the turbine (12) being mounted in the fuel circuit (CQ) so as to allow the rotational drive of the turbine (12) by the circulation of the fuel (Q),

a first heat exchanger (31), mounted upstream of the turbine (12), configured to heat the fuel (Q) in the fuel circuit (CQ) by circulation of an air flow (A),

at least one heating turbine engine (2) configured to supply the first heat exchanger (31) with an air flow (A), the heating turbine engine (2) comprising an air intake compressor (21), a combustion chamber (24) and an air exhaust turbine (22) configured to drive the air intake compressor (21), the combustion chamber (24) being fed by air taken from the air flow (A) and fuel (Q) from the fuel circuit (CQ).
System according to claim 1 comprising a second heat exchanger (32) configured to heat the fuel (Q) in the fuel circuit (CQ) by circulation of an intake air flow from the heating turbomachine (2).
System according to one of claims 1 to 2 comprising an auxiliary heat exchanger (33) configured to heat the fuel (Q) in an upstream portion of the fuel circuit (CQ) by circulation of fuel (Q) circulating downstream of the turbine (12).
System according to one of claims 1 to 3 comprising an electric generator (51) configured to be driven by the heating turbomachine (2).
System according to one of Claims 1 to 4, in which the heating turbomachine (2) is configured to implement a Brayton cycle.
System according to one of Claims 1 to 5, in which the pumping turbomachine (1) is entirely bathed in fuel (Q).
System according to one of Claims 1 to 6, in which the pumping turbomachine (1) is a hydrogen expansion turbopump.
Set of at least one cryogenic tank (RC), an aircraft turbine engine (T) and a conditioning system (SC) according to one of the preceding claims fluidically connecting the cryogenic tank (RC) and the turbine engine aircraft (T).
Assembly according to claim 8, comprising at least one fuel buffer storage capacity (7) configured to be supplied by the conditioning system (SC) and configured to supply the aircraft turbine engine (T).
Process for conditioning fuel (SC) by means of a conditioning system (SC) according to one of Claims 1 to 7 for supplying an aircraft turbine engine (T) from fuel (Q) coming from a tank cryogenic (RC), a process in which:

the pump (11) of the pumping turbomachine (1) takes fuel from the cryogenic tank (RC) and circulates it from upstream to downstream in a fuel circuit (CQ),

the fuel (Q) is heated in the fuel circuit (CQ) by the first heat exchanger (31) by circulation of an air flow (A) coming from the heating turbomachine (2),

the circulation of the fuel (Q) in the fuel circuit (CQ) rotating the turbine (12) of the pumping turbomachine (1)

the combustion chamber (24) of the heating turbine engine (2) being supplied with air taken from the air flow (A) and with fuel (Q) from the fuel circuit (CQ).